Gas turbine engine airfoil

ABSTRACT

An airfoil of a turbine engine includes pressure and suction sides that extend in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip. The airfoil has a relationship between a tangential stacking offset and a span position that is at least a third order polynomial curve that includes at least one positive and negative slope. The positive slope leans toward the suction side and the negative slope leans toward the pressure side. An initial slope starts at the 0% span position is either zero or positive. The first critical point is less than 15% span.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. application Ser. No.14/626,167 filed Feb. 19, 2015 which claims priority to U.S. ProvisionalApplication No. 61/941,698, which was filed on Feb. 19, 2014 and isincorporated herein by reference.

BACKGROUND

This disclosure relates generally to an airfoil for gas turbine engines,and more particularly to a fan or compressor blade and the relationshipbetween the blade's tangential stacking offset relative to span.

A turbine engine such as a gas turbine engine typically includes a fansection, a compressor section, a combustor section and a turbinesection. Air entering the compressor section is compressed and deliveredinto the combustor section where it is mixed with fuel and ignited togenerate a high-speed exhaust gas flow. The high-speed exhaust gas flowexpands through the turbine section to drive the compressor and the fansection. The compressor section typically includes low and high pressurecompressors, and the turbine section includes low and high pressureturbines.

The propulsive efficiency of a gas turbine engine depends on manydifferent factors, such as the design of the engine and the resultingperformance debits on the fan that propels the engine. As an example,the fan may rotate at a high rate of speed such that air passes over thefan airfoils at transonic or supersonic speeds. The fast-moving aircreates flow discontinuities or shocks that result in irreversiblepropulsive losses. Additionally, physical interaction between the fanand the air causes downstream turbulence and further losses. Althoughsome basic principles behind such losses are understood, identifying andchanging appropriate design factors to reduce such losses for a givenengine architecture has proven to be a complex and elusive task.

SUMMARY

In one exemplary embodiment, an airfoil of a turbine engine includespressure and suction sides that extend in a radial direction from a 0%span position at an inner flow path location to a 100% span position atan airfoil tip. The airfoil has a relationship between a tangentialstacking offset and a span position that is at least a third orderpolynomial curve that includes at least one positive and negative slope.The positive slope leans toward the suction side and the negative slopeleans toward the pressure side. An initial slope starts at the 0% spanposition is either zero or positive. The first critical point is lessthan 15% span.

In a further embodiment of the above, the curve has at least twocritical points.

In a further embodiment of any of the above, the airfoil extends from aroot. A zero tangential stacking offset corresponds to tangential centerof the root.

In a further embodiment of any of the above, the curve has an initialpositive slope.

In a further embodiment of any of the above, the first critical pointhas an R_(d)/Y_(d) ratio in a range of 195 to 205.

In a further embodiment of any of the above, the second critical pointis in the range of 35-45% span.

In a further embodiment of any of the above, the second critical pointhas an R_(d)/Y_(d) ratio in a range of 1950 to 2050.

In a further embodiment of any of the above, the curve does not crossthe zero tangential stacking offset.

In a further embodiment of any of the above, the curve crosses the zerotangential stacking offset once.

In a further embodiment of any of the above, a critical point in a35-45% span range has a R_(d)/Y_(d) ratio in a range of −70 to −74.

In a further embodiment of any of the above, the curve has an initialzero slope.

In a further embodiment of any of the above, the first critical point isin the range of 35-45% span.

In a further embodiment of any of the above, the first critical pointhas an R_(d)/Y_(d) ratio in a range of −108 to −116.

In a further embodiment of any of the above, the second critical pointis in the range of 85-95% span.

In a further embodiment of any of the above, the second critical pointhas an R_(d)/Y_(d) ratio in a range of 23 to 26.

In a further embodiment of any of the above, the curve crosses the zerotangential stacking offset twice.

In a further embodiment of any of the above, the airfoil is a fan bladefor a gas turbine engine.

In a further embodiment of any of the above, the fan blade is acomposite blade.

In a further embodiment of any of the above, the composite blade is aswept blade.

In another exemplary embodiment, a gas turbine engine includes acombustor section that is arranged between a compressor section and aturbine section. A fan section has an array of twenty-six or fewer fanblades. The fan section has a low fan pressure ratio of less than orequal to about 1.7. A geared architecture couples the fan section to theturbine section or the compressor section. The fan blades include aroot. An airfoil has pressure and suction sides and extends from theroot in a radial direction from a 0% span position at an inner flow pathlocation to a 100% span position at an airfoil tip. The airfoil has arelationship between a tangential stacking offset and a span positionthat is at least a third order polynomial curve that includes at leastone positive and negative slope. The positive slope leans toward thesuction side and the negative slope leans toward the pressure side. Azero tangential stacking offset corresponds to tangential center of theroot. The 0% span position is offset from the zero tangential stackingoffset. The first critical point is less than 15% span.

BRIEF DESCRIPTION OF THE DRAWINGS

The disclosure can be further understood by reference to the followingdetailed description when considered in connection with the accompanyingdrawings wherein:

FIG. 1 schematically illustrates a gas turbine engine embodiment.

FIG. 2A is a perspective view of a portion of a fan section.

FIG. 2B is a schematic cross-sectional view of the fan section.

FIG. 2C is a cross-sectional view a fan blade taken along line 2C-2C inFIG. 2B.

FIG. 3A is a schematic view of fan blade span positions.

FIG. 3B is a schematic view of a cross-section of a fan blade section ata particular span position and its tangential stacking offset.

FIG. 4 illustrates a relationship between a tangential stacking offsetand a span position for several example airfoils.

The embodiments, examples and alternatives of the preceding paragraphs,the claims, or the following description and drawings, including any oftheir various aspects or respective individual features, may be takenindependently or in any combination. Features described in connectionwith one embodiment are applicable to all embodiments, unless suchfeatures are incompatible.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmenter section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures. That is, the disclosedairfoils may be used for engine configurations such as, for example,direct fan drives, or two- or three-spool engines with a speed changemechanism coupling the fan with a compressor or a turbine sections.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis X relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a second (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 57 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axisX which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10). Inanother example, the bypass ratio between about ten (10) and abouteighteen (18). In yet another example, the bypass ratio is less thanabout nine (9), for example, about seven (7) to about nine (9). In oneexample, the geared architecture 48 is an epicyclic gear train, such asa planetary gear system or other gear system, with a gear reductionratio of greater than about 2.3 and the low pressure turbine 46 has apressure ratio that is greater than about five. In one disclosedembodiment, the engine 20 bypass ratio is greater than about ten (10:1),the fan diameter is significantly larger than that of the low pressurecompressor 44, and the low pressure turbine 46 has a pressure ratio thatis greater than about five (5:1). Low pressure turbine 46 pressure ratiois pressure measured prior to inlet of low pressure turbine 46 asrelated to the pressure at the outlet of the low pressure turbine 46prior to an exhaust nozzle. The geared architecture 48 may be anepicyclic gear train, such as a planetary gear system or other gearsystem, with a gear reduction ratio of greater than about 2.3:1. Itshould be understood, however, that the above parameters are onlyexemplary of one embodiment of a geared architecture engine and that thepresent invention is applicable to other gas turbine engines includingdirect drive turbofans.

The example gas turbine engine includes the fan 42 that comprises in onenon-limiting embodiment less than about twenty-six (26) fan blades. Inanother non-limiting embodiment, the fan section 22 includes less thanabout twenty (20) fan blades. Moreover, in one disclosed embodiment thelow pressure turbine 46 includes no more than about six (6) turbinerotors schematically indicated at 34. In another non-limiting exampleembodiment the low pressure turbine 46 includes about three (3) turbinerotors. A ratio between the number of fan blades 42 and the number oflow pressure turbine rotors is between about 3.3 and about 8.6. Theexample low pressure turbine 46 provides the driving power to rotate thefan section 22 and therefore the relationship between the number ofturbine rotors 34 in the low pressure turbine 46 and the number ofblades 42 in the fan section 22 disclose an example gas turbine engine20 with increased power transfer efficiency.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (‘TSFC’)”—is the industry standardparameter of lbm of fuel being burned divided by lbf of thrust theengine produces at that minimum point. “Low fan pressure ratio” is thepressure ratio across the fan blade alone, without a Fan Exit Guide Vane(“FEGV”) system. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.55. Inanother non-limiting embodiment the low fan pressure ratio is less thanor equal to about 1.7, for example, from about 1.2 to about 1.7. Inanother example, the low fan pressure ratio is less than or equal toabout 1.5, for example, from about 1.4 to about 1.5, and, for example,about 1.45. In another non-limiting embodiment the low fan pressureratio is from 1.1 to 1.45. “Low corrected fan tip speed” is the actualfan tip speed in ft/sec divided by an industry standard temperaturecorrection of [(Tram °R)/(518.7°R)]^(0.5). The “Low corrected fan tipspeed” as disclosed herein according to one non-limiting embodiment isless than about 1200 ft/second.

Referring to FIG. 2A-2C, the fan blade 42 is supported by a fan hub 60that is rotatable about the axis X. Each fan blade 42 includes anairfoil 64 extending in a radial span direction R from a root 62 to atip 66. A 0% span position corresponds to a section of the airfoil 64 atthe inner flow path (e.g., a platform), and a 100% span positioncorresponds to a section of the airfoil 64 at the tip 66. The airfoil 64may also be used for other sections of the engine 20, for example, thecompressor section 24 or the turbine section 28.

The root 62 is received in a correspondingly shaped slot in the fan hub60. The airfoil 64 extends radially outward of the platform, whichprovides the inner flow path. The platform may be integral with the fanblade or separately secured to the fan hub, for example. A spinner 66 issupported relative to the fan hub 60 to provide an aerodynamic innerflow path into the fan section 22.

The airfoil 64 has an exterior surface 76 providing a contour thatextends from a leading edge 68 aftward in a chord-wise direction H to atrailing edge 70, as shown in FIG. 2C. Pressure and suction sides 72, 74join one another at the leading and trailing edges 68, 70 and are spacedapart from one another in an airfoil thickness direction T. An array ofthe fan blades 42 are positioned about the axis X in a circumferentialor tangential direction Y. Any suitable number of fan blades may be usedin a given application.

The exterior surface 76 of the airfoil 64 generates lift based upon itsgeometry and directs flow along the core flow path C and the bypass flowpath B. The fan blade 42 may be constructed from a composite material,or an aluminum alloy or titanium alloy, or a combination of one or moreof these. Abrasion-resistant coatings or other protective coatings maybe applied to the fan blade 42. In one example, the fan blade 42 is acomposite that has a swept configuration.

One characteristic of fan blade performance relates to the fan blade'stangential stacking offset (Y direction) relative to a particular spanposition (R direction). Referring to FIG. 3A, span positions aschematically illustrated from 0% to 100% in 10% increments. Eachsection at a given span position is provided by a conical cut thatcorresponds to the shape of the core flow path, as shown by the largedashed lines. In the case of a fan blade with an integral platform, the0% span position corresponds to the radially innermost location wherethe airfoil meets the fillet joining the airfoil to the platform. In thecase of a fan blade without an integral platform, the 0% span positioncorresponds to the radially innermost location where the discreteplatform meets the exterior surface of the airfoil. In addition tovarying with span, tangential stacking offset varies between a hot,running condition and a cold, static (“on the bench”) condition.

The Y_(CG) corresponds to the location of the center of gravity,assuming a homogeneous material, for a particular section at a givenspan location relative to a reference point 80 in the Y direction, asshown in FIG. 3B. The reference point 80 is the tangential center of theroot, and Y_(d) corresponds to the circumferential distance from thereference point 80 to the center of gravity. A positive Y valuecorresponds to the opposite rotational direction as the hub's rotation,or toward the suction side of the airfoil. A negative Y valuecorresponds to the same rotational direction as the hub's rotation, ortoward the pressure side of the airfoil.

The tangential stacking offset Y_(d) may be expressed as a tangentialstacking offset ratio R_(d)/Y_(d), where R_(d) is the radial distancefrom hub's rotational axis X to the tip of the leading edge 68. R_(d) asdisclosed herein according to one non-limiting embodiment is about 35-37inches (0.89-0.94 meters). In another non-limiting embodiment R_(d) isabout 27-29 inches (0.69-0.74 meters). In another non-limitingembodiment R_(d) is about 39-41 inches (0.99-1.04 meters). One exampleprior art airfoil has an R_(d) of about 57-59 inches (1.45-1.50 meters).

Example relationships between the tangential stacking offset and thespan position (AVERAGE SPAN %), which is the average of the radialposition at the leading and trailing edges 68, 70. are shown in FIG. 4for several example fan blades, each represented by curves A, B and C.

The prior art curve includes a first critical point around 18% span thathas an R_(d)/Y_(d) ratio of 81-84. A second critical point around 35%span has an R_(d)/Y_(d) ratio of around 114-118, and a third criticalpoint around 95% has an R_(d)/Y_(d) ratio of around 12.7-13.1.

Each relationship is at least a third order polynomial having at leastone positive slope and one negative slope. Each inventive curve includesat least two critical points, which are represented by an asterisk. Apositive slope is where the airfoil leans toward the suction side, and anegative slope is where the airfoil leans toward the pressure side.Relationships are depicted in which the curves have an initial slopestarting at the 0% span position that is either zero or positive. Eachcurve has the 0% span position offset from the zero tangential stackingoffset, or reference point.

Referring to curve A, the curve has an initial positive slope to acritical point 90 in the range of 5-15% span. The curve begins with apositive tangential stacking offset. At the 5-15% span location, thetangential stacking offset ratio R_(d)/Y_(d) is in the range of 195 to205. The curve has a negative slope from the critical point 90 to ancritical point 92, which is in the range of 35-45% span. At the 35-45%span location, the tangential stacking offset ratio R_(d)/Y_(d) is inthe range of 1950 to 2050. At the 100% span location, the tangentialstacking offset ratio R_(d)/Y_(d) is in the range of 13 to 15. The curvedoes not cross the zero tangential stacking offset.

Referring to curve B, the curve has an initial positive slope to acritical point 94 in the range of 5-15% span. The curve begins with anegative tangential stacking offset. At the 5-15% span location, thetangential stacking offset ratio R_(d)/Y_(d) is in the range of −175 to−185. The curve has a negative slope from the critical point 94 to acritical point 96, which is in the range of 35-45% span. At the 35-45%span location, the tangential stacking offset ratio R_(d)/Y_(d) is inthe range of −70 to −74. The curve crosses the zero tangential stackingoffset at location 98 in the range of 60-70% span. At the 100% spanlocation, the tangential stacking offset ratio R_(d)/Y_(d) is in therange of 19 to 21.

Referring to curve C, the curve has an initial zero slope and beginswith a positive tangential stacking offset. The curve has a negativeslope that crosses the zero tangential stacking offset at location 104after a critical point in the range of 5-15% and has a critical point100 in the range of 35-45% span. At the 5-15% span location, thetangential stacking offset ratio R_(d)/Y_(d) is in the range of −540 to−580. At the 35-45% span location, the tangential stacking offset ratioR_(d)/Y_(d) is in the range of −108 to −116. The curve has a positiveslope from the critical point 100 to a critical point 102, which is inthe range of 85-95% span. The curve crosses the zero tangential stackingoffset in the range of 50-60% span. At the 85-95% span location, thetangential stacking offset ratio R_(d)/Y_(d) is in the range of 23 to26. At the 100% span location, the tangential stacking offset ratioR_(d)/Y_(d) is in the range of 27 to 29.

The tangential stacking offset in a hot, running condition along thespan of the airfoils 64 relate to the contour of the airfoil and providenecessary fan operation in cruise at the lower, preferential speedsenabled by the geared architecture 48 in order to enhance aerodynamicfunctionality and thermal efficiency. As used herein, the hot, runningcondition is the condition during cruise of the gas turbine engine 20.For example, the tangential stacking offsets in the hot, runningcondition can be determined in a known manner using numerical analysis,such as finite element analysis.

It should also be understood that although a particular componentarrangement is disclosed in the illustrated embodiment, otherarrangements will benefit herefrom. Although particular step sequencesare shown, described, and claimed, it should be understood that stepsmay be performed in any order, separated or combined unless otherwiseindicated and will still benefit from the present invention.

Although the different examples have specific components shown in theillustrations, embodiments of this invention are not limited to thoseparticular combinations. It is possible to use some of the components orfeatures from one of the examples in combination with features orcomponents from another one of the examples.

Although an example embodiment has been disclosed, a worker of ordinaryskill in this art would recognize that certain modifications would comewithin the scope of the claims. For that reason, the following claimsshould be studied to determine their true scope and content.

What is claimed is:
 1. An airfoil of a turbine engine comprising:pressure and suction sides extending in a radial direction from a 0%span position at an inner flow path location to a 100% span position atan airfoil tip, wherein the airfoil has a relationship between atangential stacking offset and a span position that is at least a thirdorder polynomial curve that includes at least one positive and negativeslope, the positive slope leans toward the suction side and the negativeslope leans toward the pressure side, wherein an initial slope startingat the 0% span position is either zero or positive, wherein the firstcritical point is less than 15% span.
 2. The airfoil according to claim1, wherein the curve has at least two critical points.
 3. The airfoilaccording to claim 2, wherein the airfoil extends from a root, and azero tangential stacking offset corresponds to tangential center of theroot.
 4. The airfoil according to claim 2, wherein the curve has aninitial positive slope.
 5. The airfoil according to claim 4, wherein thefirst critical point has an R_(d)/Y_(d) ratio in a range of 195 to 205.6. The airfoil according to claim 4, wherein the second critical pointis in the range of 35-45% span.
 7. The airfoil according to claim 6,wherein the second critical point has an R_(d)/Y_(d) ratio in a range of1950 to
 2050. 8. The airfoil according to claim 4, wherein the curvedoes not cross the zero tangential stacking offset.
 9. The airfoilaccording to claim 4, wherein the curve crosses the zero tangentialstacking offset once.
 10. The airfoil according to claim 9, wherein acritical point in a 35-45% span range has a R_(d)/Y_(d) ratio in a rangeof −70 to −74.
 11. The airfoil according to claim 2, wherein the curvehas an initial zero slope.
 12. The airfoil according to claim 11,wherein the first critical point is in the range of 35-45% span.
 13. Theairfoil according to claim 12, wherein the first critical point has anR_(d)/Y_(d) ratio in a range of −108 to −116.
 14. The airfoil accordingto claim 11, wherein the second critical point is in the range of 85-95%span.
 15. The airfoil according to claim 14, wherein the second criticalpoint has an R_(d)/Y_(d) ratio in a range of 23 to
 26. 16. The airfoilaccording to claim 11, wherein the curve crosses the zero tangentialstacking offset twice.
 17. The airfoil according to claim 1, wherein theairfoil is a fan blade for a gas turbine engine.
 18. The airfoilaccording to claim 17, wherein the fan blade is a composite blade. 19.The airfoil according to claim 18, wherein the composite blade is aswept blade.
 20. A gas turbine engine comprising: a combustor sectionarranged between a compressor section and a turbine section; a fansection having an array of twenty-six or fewer fan blades, wherein thefan section has a low fan pressure ratio of less than or equal to about1.7; a geared architecture coupling the fan section to the turbinesection or the compressor section; and wherein the fan blades include aroot, and an airfoil having pressure and suction sides extending fromthe root in a radial direction from a 0% span position at an inner flowpath location to a 100% span position at an airfoil tip, wherein theairfoil has a relationship between a tangential stacking offset and aspan position that is at least a third order polynomial curve thatincludes at least one positive and negative slope, the positive slopeleans toward the suction side and the negative slope leans toward thepressure side, wherein a zero tangential stacking offset corresponds totangential center of the root, and the 0% span position is offset fromthe zero tangential stacking offset, wherein the first critical point isless than 15% span.